Most relatively small missiles in use today are propelled by solid fuel rockets as opposed to, for example, turbine engines. The selection of a solid fuel rocket as a propulsion device has been largely dictated by two factors. First, in many instances, a turbine engine cannot be fabricated sufficiently economically as to compete with a solid fuel rocket engine. Secondly, in small sized missiles, i.e., those having relatively small diameters on the order of about four inches, it has heretofore been quite difficult to manufacture an efficient turbine engine. The difficulty lies in the fact that the turbine engine must fit within the four inch envelope required of the propulsion unit for such a missile. Unfortunately, the use of solid fuel rocket engines has had consequences that are not entirely desirable in many applications.
Specifically, the use of solid fuel rocket engines results in the loss of some degree of control of the missile flight path or trajectory. In contrast, control is far greater with gas turbine engines whose output can readily be varied. Further, even if the gas turbine engine operates relatively inefficiently, the use of such an engine would greatly extend the range of the missile.
As will be appreciated, one of the difficulties in economically producing small diameter gas turbine engines resides in the labor intensive nature of the manufacture of the combustor. Furthermore, as combustor sizes shrink so as to fit within some desired envelope, the difficulty in achieving efficient combustion of fuel rises significantly. In particular, as the size or volume of a combustor is reduced, there may be insufficient volume to allow the fuel to first be vaporized completely, burned efficiently, and then mixed uniformly.
In order to overcome the foregoing, a unique low cost annular combustor was developed as disclosed in commonly owned U.S. Pat. No. 4,794,754, issued Jan. 3, 1989. This annular combustor has proven to be well suited for its intended purpose, but it was desired to attempt to achieve greater performance characteristics with a higher turbine inlet temperature while meeting the necessary size constraints and achieving the goal of simplicity coupled with ultralow cost for throw away missile applications. For this purpose, it was recognized that a new approach would be required to completely reach the necessary parameters of operation.
In order to achieve these objectives, a combustor having enhanced turbine nozzle cooling was developed as disclosed in commonly owned U.S. Pat. No. 4,825,640, issued May 2, 1989. And as a further improvement in this field, an inexpensive annular combustor promoting both full and efficient vaporization and combustion along with thorough mixing for uniform exit temperatures was developed. More specifically, such a combustor is disclosed in commonly owned and co-pending patent application U.S. Ser. No. 455,588, filed Dec. 22, 1989.
While all of these represent significant advancements in the field, it is believed that still further enhanced operating parameters are capable of being achieved. Accordingly, the present invention is directed to still more fully overcoming the foregoing problems while achieving this objective to the fullest extent possible.